This application claims the priority of German Patent Document DE 101 26 926.9, filed on Jun. 1, 2001, the disclosure of which is expressly incorporated by reference herein.
This invention relates to a combustion chamber, in particular for a rocket drive, exhibiting at least one jacket made of a composite material with a ceramic matrix, wherein the composite material contains a fibrous structure made of carbon-containing fibers, as well as a process for manufacturing a combustion chamber, in particular for a rocket drive, wherein at least one jacket is made of a composite material with a ceramic matrix, wherein the composite material is made on the basis of a fibrous structure made of carbon-containing fibers and wherein silicon is fed to the fibrous structure in order to form a silicon carbide matrix. The invention is applicable, in principle, not only to the special application of rocket drive, engineering but also to other fields, such as aircraft engineering, and, in principle, also to other combustion chambers and furnaces.
The prior art discloses combustion chambers, in particular for rocket drives, where composite materials with a ceramic matrix are used. The disclosure of DE 198 58 197 describes a drive with a combustion chamber, provided with an internal jacket and insulation made of carbon fiber-reinforced silicon carbide (C/SiC). DE 197 30 674 discloses a combustion chamber, in particular for a rocket drive, provided with an internal jacket made of a fiber-reinforced ceramic material or graphite and an external jacket made of a fiber-reinforced ceramic material. DE 37 34 100 describes a combustion chamber for an aircraft, wherein a porous, woven internal wall made of silicon carbide fibers and a woven external wall made of silicon carbide fibers are provided.
A significant problem with the prior art construction method of the combustion chambers with at least one jacket made of a composite material with a ceramic matrix is that, for the ceramic matrix composites (CMC) materials or comparable composite materials used at that time, two-dimensional (2D) structures, in particular 2D C/SiC or 2D carbon fiber reinforced carbon (C/C), are used. These structures exhibit only negligible interlaminar shear strength (ILS) in relation to the shear forces, as can be the case especially in the thrust direction of a rocket drive. For example, U.S. Pat. No. 6,197,411 describes such two-dimensional structures for composite materials with a metal matrix. The stress conditions that can occur in the combustion chamber and that are induced by corresponding thermal loads can result in delamination between individual layers of the CMC structures, a factor that can impair the operability of the combustion chamber and can result in the failure of the combustion chamber.
For special applications of the rocket drives, mechanical stresses such as vibration, bending strain, acoustical loads, that can occur especially at launch time, must be taken into account. To this end, it is necessary that the structure of the drive exhibit adequate ability to absorb any loads and at the same time demonstrate adequate elastic deformation or ductility. The ceramic external structures, known from the prior art, do in fact exhibit high rigidity, but also poor expansion and deformation properties. In addition, the external structure must be gas-tight. The manufacture of gas-tight fiber-reinforced composite ceramics, for example, C/SiC or C/C, can be realized only at a very high cost according to the prior art methods.
Therefore, one object of the present invention is to provide a combustion chamber and a method for manufacturing a combustion chamber, in particular for a rocket drive, which overcome the drawbacks of the prior art.
According to one embodiment, the invention comprises a combustion chamber, in particular for a rocket drive, exhibiting at least one jacket made of a composite material with a ceramic matrix, wherein the composite material contains a fibrous structure made of carbon-containing fibers. In addition to applications in rocket drive engineering, such a combustion chamber can be used for combustion chambers of aircraft technology or other combustion chambers or furnaces. According to the invention, the fibrous structure consists of layers of fibers, forming a three-dimensional matrix. Thus, there is no longer a two-dimensional configuration of the individual fibers or the layers of fibers, but rather the fibers or the layers of fibers are arranged in a defined way in the form of a three-dimensional matrix. In this respect, the fibers or layers of fibers are suitably interconnected in order to guarantee a defined matrix structure. Such a structure significantly improves the stability of the composite material with respect to shear stresses. In addition, the goal with such a defined three-dimensional configuration of fibers is to improve and optimize the adjustability of the thermal conductivity of the composite material.
In particular, it can be provided that the fibrous structure be constructed from first, second and third layers of fibers, wherein the fibers of the first layers extend in a first direction in space; the fibers of the second layers extend in a second direction in space; and the fibers of the third layers extend in a third direction in space; and wherein the individual layers penetrate each other at least partially. Thus, the individual layers can comprise, for example, fibers or bundles of fibers that are arranged parallel to each other, wherein the fibers or bundles of fibers of each layer are separated from each other so that fibers or bundles of fibers of another layer, extending in another direction in space, can be disposed in the resulting spaces. This feature offers the possibility of mutual interpenetration of the individual layers.
Furthermore, it can be provided that the individual layers of the fibrous structure be interconnected by means of textile technology. For example, the individual layers can be interwoven or sewn together. This feature represents a simple and effective way to interconnect individual layers of fibers or bundles of fibers.
In the case of previous combustion chambers, it was provided that silicon carbide fibers be used. However, they have a drawback because they remain stable only up to temperatures of approximately 1,200° C. For higher temperature ranges, for example at 1,500° C. or above, as can occur in a rocket drive, such fibers are inappropriate or they are appropriate only with the provision of additional protective measures. Therefore, for applications at higher operating temperatures exceeding 1,500° C., the fibrous structure made of carbon fibers is preferred.
A composite material that contains silicon carbide can be provided as a special composite material. This silicon carbide can also be formed, at least to some degree, through a reaction of silicon with the fibrous structure.
For applications of the combustion chamber under high loads, as particularly in a rocket drive, or in a correspondingly low load-bearing structure of the jacket made of composite material, it is provided that the jacket made of composite material be enveloped by a load-bearing external jacket. This external jacket serves to support the composite material jacket, which then forms a corresponding internal jacket. To achieve adequate flexibility precisely for applications in a rocket drive, said external jacket is made preferably of a metal material, such as nickel, copper, or a nickel- and/or copper-based alloy, or a suitable steel.
If, as a function of the chosen type of materials and structures of the internal jacket made of composite material and of the external jacket, the result is that these two jackets have a significantly different thermal expansion coefficient, then an intermediate layer, whose thermal expansion coefficient is between that of the external jacket and that of the composite material jacket, can be provided to prevent stresses or even cracks between the external jacket and the composite material jacket. For example, a composite material with a metal matrix can be provided as the intermediate layer, but any other suitable material can also be provided. The intermediate layer of metal-matrix composite material can be affixed on the composite material jacket. To achieve a balance between the thermal expansion coefficients that is as simple and effective as possible, it is preferably provided that the metal matrix contains the same metal material as the external jacket. The external jacket can be affixed on the intermediate layer.
The invention also comprises a process for manufacturing a combustion chamber, in particular for a rocket drive, wherein at least one jacket is made of a composite material with a ceramic matrix, wherein the composite material is based on a fibrous structure made of carbon-containing fibers, and wherein the silicon is fed to the fibrous structure for the purpose of forming a silicon carbide matrix. The invention provides that the fibrous structure be produced as a three-dimensional matrix comprising layers of fibers. Thus, the aforementioned advantages of increased stability, with respect to the shear stresses, and better adjustability of the thermal conductivity, are achieved in a simple and defined manner.
Channel-shaped cooling structures can be provided in the area of the fibrous structure. These cooling structures can be produced in many ways. The layers of the fibrous structure can be arranged in such a manner that channel-shaped spaces remain on the surface of the fibrous structure and/or in the fibrous structure. In this way, the cooling structures are already defined by way of the fibrous structure itself, a feature that has the advantage that the fibers are not destroyed by subsequent treatment steps. It can also be provided that channel structures be not worked into the surface of the composite material by way of mechanical treatment operations until after the composite material has been fabricated. This manufacturing method is relatively simple to carry out. However, channel-shaped contracting bodies, which decompose or dissolve in the course of manufacturing the composite material or thereafter and open the corresponding channels in the composite material, can also be placed, for example, in the fibrous structure within the framework of manufacturing the fiber matrix.
An additional seal and load-bearing reinforcement of the channel structures can be produced in that at least those surface areas of the composite material that exhibit the channel structures are coated with a metal material. Metal-lined channel structures are then produced that better fulfill the requirements imposed on the coolant, flowing through the channels, at least for specific areas of application.
If channel structures are provided on the fibrous structure and other structures are supposed to be produced on this fibrous structure, then it can be provided that channel-shaped contracting bodies be disposed on the surface of the fibrous structure, precisely in the channel structures present there. Then they can be decomposed or dissolved, as described above, in the course of manufacturing the composite material or thereafter and open the corresponding channels in the composite material.
For applications in rocket drives, a load-bearing external jacket can be affixed on the jacket made of composite material. For the aforementioned reasons, it can be provided that an external jacket made of a metal material be affixed. Such an external jacket can be affixed by way of electroplating, soldering, welding techniques, or other known methods.
There are, in principle, various alternatives for composite materials with a metal matrix that can be used for the intermediate layer. Their manufacture is known to some degree. See, for example, U.S. Pat. No. 6,197,411. However, an improved method for manufacturing such a composite material, in particular for forming an intermediate layer between the internal jacket and the external jacket of a combustion chamber, would be an improvement in the art, and such an improvement is provided by the present invention. First, a fibrous structure is affixed on the composite material jacket, and then a metal material is deposited on the fibrous structure with simultaneous infiltration of the fibrous structure with the metal material. Thus, in one working step, a composite material with metal matrix can be produced as the intermediate layer, and an external jacket can be produced on the intermediate layer. The metal material can be deposited in any suitable form, for example, from a liquid or gaseous phase. However, the metal material can also be deposited by way of an electroplating process. In particular nickel or copper or a nickel- and/or copper-containing alloy can be used here as the metal material.
The invention also comprises a process for manufacturing an intermediate layer between an internal jacket and an external jacket of a combustion chamber, in particular a rocket drive, wherein at least one part of the structure is produced from a composite material on the basis of a fibrous structure made of carbon-containing fibers. First, a fibrous structure made of carbon-containing fibers is affixed on the internal jacket, and then a metal material is deposited on the fibrous structure with simultaneous infiltration of the fibrous structure with the metal material. The metal material can be deposited in any suitable form, for example, in a liquid or gas phase. However, the metal material can also be deposited preferably by way of an electroplating process. In particular, nickel or copper or a nickel- and/or copper-containing alloy can be used here as the metal material.
Other objects, advantages and novel features of the present invention will become apparent from the following detailed description of the invention when considered in conjunction with the accompanying drawings.